The present invention relates generally to a cooling system for the nozzle segments of a gas turbine and particularly relates to a cooling system for cooling the adjoining edges of inner and outer platforms of adjacent nozzle segments arranged in an annular array about the axis of the turbine.
In gas turbines, annular arrays of nozzles are disposed in the hot gas path for turning and accelerating the gas flow for optimum performance of the buckets. In the first stage of a turbine, for example, there are a plurality of circumferentially spaced nozzle vanes which extend generally radially between inner and outer annular bands which serve to confine the gas flow to an annular configuration as the gas flows through the multiple stages of the turbine. A plurality of circumferentially spaced buckets mounted on the turbine rotor lie axially downstream of the annular array of nozzles and form a turbine stage with the nozzles. The nozzles, for example, of the first stage of the turbine, are typically provided in nozzle segments. Each nozzle segment includes an inner platform and an outer platform and at least one vane extending between the platforms. The nozzle segments are arranged in circumferential registration with one another. Particularly, the inner and outer platforms of each nozzle segment lie in circumferential registration with the inner and outer platforms of adjacent segments, respectively. In this arrangement, gaps are formed between adjoining segments  along the platform edges. Prior nozzle platform edges have been uncooled, cooled by film cooling from adjacent nozzle segments or cooled by long holes that run from a large impingement cavity in the nozzle segment to the gaps between the nozzle segments. Film cooling from an adjacent nozzle to cool the platform edge, however, causes a debiting of the cooling effectiveness when the cooling film crosses the nozzle intersegment gap. When long holes running from an impingement cavity are utilized, the convective cooling of the edge by the holes is discrete rather than continuous and, therefore, less efficient.
Certain prior nozzle designs have adjacent platform edges configured such that the nozzle intersegment gaps are aligned parallel to the hot gas flow vector. Perfect alignment of the adjoining edges of the nozzle segments, however, is difficult to achieve and maintain as a result of manufacturing and thermomechanical problems. It will be appreciated that the core flow boundary layers of the hot gas along the platform surfaces may be tripped if the intersegment gap is not aligned with the flow direction. A boundary layer trip at the adjoining edges of the platforms results in a spike in heat transfer near the edge of the platform and also results in a debit to the cooling effectiveness of any film cooling medium that crosses the gap.
Notwithstanding the desirability of aligning the inner segment gaps parallel to the flow vector, it is beneficial for other reasons to provide nozzle platform edges which extend generally parallel to the axis of the rotor. This enables removal of the nozzles without  removal of the top half of the turbine shell, resulting in less expensive and more flexible maintenance. Consequently, the intersegment gap is not aligned with the core flow downstream of the vane. Such design is more sensitive to any platform deformations that would cause a mismatch between the platform edges of adjacent nozzle segments and cause the core flow to “see” a facing step. Thus, the edges of nozzles segment platforms which extend generally parallel to the turbine axis are subject to severe thermal distress due to boundary layer trip. Accordingly, it has been found desirable to provide a cooling system which would minimize or eliminate the foregoing problems associated with cooling edges of nozzle segments wherein the edges lie generally parallel to the turbine axis.